Engineering topics

Bouslog, Stanley A.

Publications and source records attributed to Bouslog, Stanley A..

9 records

NASA NTRS2017

Characterization of a 50kW Inductively Coupled Plasma Torch for Testing of Ablative Thermal Protection Materials

With the development of new manned spaceflight capabilities including NASA's Orion capsule and the Space-X Dragon capsule, there is a renewed importance of understanding the dynamics of ablative thermal protection systems. To this end, a new inductively coupled plasma torch facility is being developed at UT-Austin. The torch operates on argon and/or air at plasma powers up to 50 kW. In the present configuration the flow issues from a low-speed subsonic nozzle and the hot plume is characterized using slug calorimetry and emission spectroscopy. Preliminary measurements using emission spectroscopy have indicated that the torch is capable of producing an air plasma with a temperature between 6,000 K and 8,000 K depending on the power and flow settings and an argon plasma with a temperature of approximately 12,000 K. The operation envelope was measured, and heat flux measured for every point within the envelope using both a slug calorimeter and a Gardon gauge heat flux sensor. The torch was found to induce a stagnation point heat flux of between 90 and 225 W/sq cm.

Greene, Benton R.↗

NASA NTRS1998

X-33 Metallic TPS Tests in NASA-LaRC High Temperature Tunnel

Conclusions The first series of metallic TPS tests in the NASA-LaRC Mach 7 High Temperature Tunnel has been completed. Additional testing is 'in progress and shall provide data for off-design configurations for the metallic TPS. The available data are being analyzed and being used to correlate analytical models to be used for X-33 flight design analysis. The final paper shan present additional data from these tests and comparisons between the data and analytical predictions.

Bouslog, Stanley A.↗

NASA NTRS1997

Isolated Roughness Induced Boundary-Layer Transition: Shuttle Orbiter Ground Tests and Flight Experience

Recent Orbiter wind tunnel data and flight data have been used to evaluate boundary-layer transition induced by discrete-roughness elements on the Orbiter windward surface. Orbiter flow field calculations have been used to compute transition parameters and disturbance parameters for correlating the results and comparing the trends. Existing transition correlations have been modified and applied to the Orbiter. These correlations provide a means to predict transition on the Orbiter given a known isolated roughness element. Furthermore, these data and data reduction methods provide information and guidance for the prediction of transition due to &screte-roughness elements on future winged reentry vehicles.

Bouslog, Stanley A.↗

NASA NTRS1993

Two-layer convective heating prediction procedures and sensitivities for blunt body reentry vehicles

This paper provides a description of procedures typically used to predict convective heating rates to hypersonic reentry vehicles using the two-layer method. These procedures were used to compute the pitch-plane heating distributions to the Apollo geometry for a wind tunnel test case and for three flight cases. Both simple engineering methods and coupled inviscid/boundary layer solutions were used to predict the heating rates. The sensitivity of the heating results in the choice of metrics, pressure distributions, boundary layer edge conditions, and wall catalycity used in the heating analysis were evaluated. Streamline metrics, pressure distributions, and boundary layer edge properties were defined from perfect gas (wind tunnel case) and chemical equilibrium and nonequilibrium (flight cases) inviscid flow-field solutions. The results of this study indicated that the use of CFD-derived metrics and pressures provided better predictions of heating when compared to wind tunnel test data. The study also showed that modeling entropy layer swallowing and ionization had little effect on the heating predictions.

Bouslog, Stanley A.↗

NASA NTRS1993

Thermal Protection System design studies for lunar crew module

The results of a study to predict aeroheating and Thermal Protection System (TPS) requirements for manned entry vehicles returning to Earth from the moon are presented. The effects of vehicle size and lunar-return strategies on the aerothermodynamic environment and TPS design were assessed. Study guidelines were based on an Apollo Command Module (CM) configuration and lunar return strategies included direct entry and aerocapture followed by Low Earth Orbit entry (LEO). Convective heating was obtained by the boundary layer integral matrix procedure (BLIMP) code, and radiative heating was computed with the QRAD program. The AESOP-STAB code and the AESOP-THERM code were used for TPS analysis for ablating materials and nonablating materials respectively. Results indicated that there was an optimum size for minimum heating and that direct entry had higher heating rates than aerocapture. Aerocapture resulted in higher heat loads and TPS weight. The TPS weight factor was 6-8 percent for all lunar return strategies, with the TPS weight being about 50 percent less than that of the original Apollo CM vehicle.

Williams, S. D.↗

NASA NTRS1993

Emittance measurements of Space Shuttle Orbiter Reinforced Carbon-Carbon

The spectral and total normal emittance of the Reinforced Carbon-Carbon (RCC) used on Space Shuttle Orbiter nose-cap and wing leading edges has been measured at room temperature and at surface temperatures of 1200 to 2100 K. These measurements were made on virgin and two flown RCC samples. Room temperature directional emittance data were also obtained and were used to determine the total hemispherical emittance of RCC as a function of temperature. Results of the total normal emittance for the virgin samples showed good agreement with the current RCC emittance design curve; however, the data from the flown samples showed an increase in the emittance at high temperature possibly due to exposure from flight environments.

Caram, Jose M.↗

NASA NTRS1993

Reentry analysis for low Earth orbiting spacecraft

As a result of recent National Aeronautics and Space Administration (NASA) Management Instruction (NMI), NASA spacecraft programs must limit orbital debris by design and/or by operational procedures. To fulfill this requirement, spacecraft may be required to be removed from their operational orbit after mission completion. Spacecraft disposal by atmospheric reentry is a means to accomplish this task. To assess the risk to man, an analysis must be done to determine which parts of the spacecraft are likely to survive a reentry of the Earth's atmosphere and where those parts will land. These issues are currently being examined for the Earth Observing System (EOS-AM1). The Johnson Space Center (JSC) Aeroscience Branch, supported by the Lockheed Engineering and Sciences Co., has developed a tool which permits the analysis of the thermal effects of reentry on individual spacecraft components to determine which components are expected to survive reentry. This paper presents an examination of the burnup and reentry of EOS-AM1 and describes a method for other spacecraft to use in analyzing similar reentry issues.

Newman, Lauri Kraft↗

NASA NTRS1992

Emittance measurements of Space Shuttle orbiter reinforced carbon-carbon

The spectral and total normal emittance of the Reinforced Carbon-Carbon (RCC) used on Space Shuttle nose cap and wing leading edges has been measured at room temperature and at surface temperatures of 1200 to 2100 K. These measurements were made on virgin and two flown RCC samples. Room temperature directional emittance data were also obtained and were used to determine the total hemispherical emittance of RCC as a function of temperature. Results of the total normal emittance for the virgin samples showed good agreement with the current RCC emittance design curve; however, the data from the flown samples showed an increase in the emittance at high temperature possibly due to exposure from flight environments.

Caram, Jose M.↗

NASA NTRS1992

Emittance measurements of RCG coated Shuttle tiles

The spectral and total normal emittance of the Reaction Cured Glass (RCG) coating used on Shuttle tiles has been measured for surface temperatures of 300 to 1905 K. These measurements were made on two virgin and two flown Shuttle tile samples. Room temperature directional emittance data were also obtained and used to determine the total hemispherical emittance of RCG as a function of temperature. The data obtained from this calculation indicate that the total hemispherical emittance decreases from a room temperature value of 0.83 to a value of 0.76 at 1905 K. The flown Shuttle tiles exhibited a change in the spectral distribution of emittance compared to that of the virgin tile, but no significant trends in the total emittance from a virgin to a flown tile could be established.

Bouslog, Stanley A.↗